(a) Field of the Invention
The present invention pertains to spacecraft attitude sensors and, more particularly, to an architecture for interfacing an attitude sensor to a processing unit.
(b) Description of Related Art
The term attitude is used to describe the orientation of an object with respect to a reference orientation. Attitude is of particular interest in satellite operations. For example, if a satellite is to be used in a communications application, it is necessary that the satellite be oriented in the proper direction to receive and/or transmit relevant information for the communication link. The attitude of a satellite is determined by computations based on the output of attitude sensors located on the satellite. Attitude sensors monitor the position of reference objects, which may include the sun, the earth, constellations, or beacons. For example, a spacecraft may have two sensors, one monitoring the sun and one monitoring the earth. It is well known that given position information from two reference objects (e.g., the sun, the earth, stars, or beacons), it is possible to determine spacecraft attitude.
Commonly, sun sensors include electronics for processing signals representative of solar position. Sun sensor electronics require power connections, telemetry and command (T&C) connections, and connections to transfer the output of the sensor electronics to the spacecraft computer. Typically, each sensor is cross-strapped to redundant spacecraft computers, power buses, and T&C subsystems. Redundant sensors are cross-strapped because of the relatively low reliability of the signal processing electronics in comparison to the other components of the sensor.
In known sensor configurations, a dedicated sensor electronics box may be disposed remote from the sun sensor head. This configuration does not simplify the interface configuration needed to link the sensor electronics box to a power supply and the telemetry system of the spacecraft. A conventional approach is to feed current signals from the sun sensor to a dedicated sensor electronics box that converts the current signals to attitude measurements which are output to the spacecraft computer as digital data. As previously mentioned, power and telemetry signals must also be fed from the spacecraft to the dedicated sensor electronics.
In order to eliminate the need for sensor electronics altogether, some known sun sensor configurations process sensor output signals within the spacecraft computer. However, these configurations still rely on conventional sensor readings which tend to limit overall system precision and accuracy.